Saturday, December 10, 2016

Update 6-19-18: This article seems to be drawing a lot of readership lately. I hope readers are finding it useful. The booster issues discussed in the comments and replies to this article are thoroughly covered in my book. That book, submitted to AIAA for publication, is "still out for review" every time I have inquired about its status, for the last several months. You might contact AIAA and inquire about it yourself, if you are interested in obtaining a copy. The title as submitted is "A Practical Guide to Ramjet Propulsion". They might move more quickly if they thought there was real demand.------------------------------------------------------Update 11-28-17: The last draft materials (that 4th appendix and a revised table of contents) have been submitted to AIAA. I haven't yet heard back on a publishing decision. ------------------------------------------------------Update 9-19-17: Application materials and most of the book went to AIAA today. Their timelines say weeks for approval, months for actual publication. ---------------------------------------------------------------------

Update 5-17-17: My ramjet book is nearly completed. I have drafted 22 of 22 chapters, and 3 of 4 appendices. I am in process of applying to publish this work as a book in AIAA's "education series".

---------------------------------------------------------------------

Many folks have looked at my Ramjet Cycles Analysis posting on this site, dated 12-21-2012. There is a whole lot more to engineering ramjet propulsion than just cycle analysis. The following is but a primer on this very large topic. It is not comprehensive, nor does it have the real "how-to". But is does introduce the real operative concepts. For the details and "how-to", you'll have to wait for my book.

To see the figures enlarged, click on one, and you can see them all. The "escape key" gets you out of the enlarged figures and back to the normal view.

There are two different
speed ranges for conventional subsonic combustion ramjets, with two
different sets of appropriate design features. There’s no point trying to
use either one outside its appropriate speed range.

Low Speed Ramjets

The low speed range
extends from very high subsonic to about Mach 2, no more than about
2.5, and such designs have a simple
pitot/normal shock inlet, and a
convergent-only nozzle that is not always choked. That nozzle in a
well-designed system begins to choke in the vicinity of Mach 1 to 1.1,
which limits combustor flow speeds to values compatible with successful
flameholding (no more than about Mach 0.42).
These have been off-the-shelf ready-to-apply liquid-fueled ramjet technology
since WW2. Example: Red-Head / Roadrunner, Gorgon-IV.
Rather volatile fuels like gasoline or the wide-cut fuel JP-4 are
required.

High-Speed Ramjets

The high-speed designs
extend from just under Mach 2 to speeds above Mach 4, to maybe Mach
6. These have supersonic inlets with external compression features that don't
work at all below about Mach 1.7, give
or take ~ 0.2. The nozzle is an always-choked convergent-divergent
design with a very modest expansion ratio (near 1.5-1.7). In a
properly-sized system, combustor flow speeds are flameholding-limited to
the same values as in a low-speed design, but pressures are higher
because the incoming inlet pressures and air massflow are higher, so
performance potential is higher than the low speed designs. These
have been off-the-shelf, ready-to-apply liquid-fueled
ramjet technology since the 1950's. Examples: Bomarc,
Talos, SA-4, SA-6,
Bloodhound, Navajo.

Superior flame
stabilization techniques, solid gas
generator-fed fuel options, and better
ways to add the needed booster rockets got added in the 1960's, and have
been off-the-shelf ready-to-apply technologies since about 1970. Examples:
SA-6, ALVRJ, ASALM-PTV,
Kh-31, Kh-41, Kh-61,
and Meteor. The prime innovation
was the integral booster. There are also
ejectable booster nozzles, nozzle-less
boosters, and inlet port cover design
approaches, all associated with those
integral boosters. The other booster
designs are stage-off (Talos) and carry-along (Bomarc). Slightly-less volatile kerosene and
kerosene-like fuels may be used in these designs, and solid gas generator-fed fuels are
feasible. There are severe geometry
restrictions with those solid fuels that require flameholding, less with the hypergolic solids.

Characteristics of Ramjets

Neither type (low- or
high-speed) operates at constant fuel flow rate, unless you only fly at only one
speed and altitude. Inlet captured airflow varies very strongly with
speed of flight and with ambient air pressure, which at high altitudes is
very low indeed. The variations are not linear, reflecting
both atmospheric variations and actual inlet hardware characteristics, as
well as engine operating parameters like fuel flow.

What
you want is operation at a constant fuel/air ratio, thus your fuel flow varies exactly as
wildly as your airflow capture does. A thrust-intense accelerator mixture
might be 10% rich. Higher fuel economy performance is available
leaned-back in steady cruise at around 10-15% lean. This leaned value
varies a lot from design to design. The rich value is an almost universal
“knee-in-the-curve” item.

For both kinds of
design, frontal thrust density (thrust per unit cross section area)
varies more-or-less in proportion to operating combustor pressure, which
at high altitudes is very low indeed. Weight doesnot

reduce with reducing
air pressure, so there quickly comes a point where you cannot generate
enough thrust to actually fly level, much less climb or accelerate.
Very few designs can successfully operate much above about
80-100,000 feet. There have been a very few exceptions, but nothing above 125,000 feet.

“Scramjets”

Supersonic combustion ramjets (“scramjets”) operate from about Mach 4 on up, albeit
at lower fuel economy performance than high speed ramjets have in their speed
range. They also have the same low frontal thrust density limitations
as the conventional ramjet. They must also variably-meter fuel at
constant mixture ratio with a highly-variable captured airflow, but
beyond that, there is no resemblance to conventional ramjet. Try to
run one below Mach 4, and it
explodes, according to the test data.

Scramjets share only a
few inlet features with high speed conventional ramjets. The design
analysis techniques, and the basic hardware components for the combustor
and nozzle, are entirely different from what is used for conventional
ramjet. Scramjet cycle analysis is very effort-intensive and
usually based on computer fluid dynamic-type analysis. Conditions are so far outside
what these codes were developed for, your analyses must be validated
by test before you can trust them. That has been rather
unreliable. Scramjets require a long
constant-Mach isolator duct, while ramjets
require a terminal shock in a divergent subsonic diffuser. These are fundamentally-incompatible
geometry requirements. Scramjets
just explode without the isolator duct.
Ramjets cannot function without the subsonic diffuser. Simple as that.

None of that scramjet
“sophisticated non-ideal-gas” analysis is needed for conventional ramjet,
coming as it does from the pencil/paper/slide rule days, and
operating in regimes where the ideal gas models still apply. While ramjet
can be done with scramjet analysis tools,it is hardly worth the effort and cost to do so.It is simply far more effective and efficient
to “do it the old way”. There is no way around that little fact of life.

The recent test flight
vehicles X-43A and X-51A did demonstrate successful, but very
experimental scramjet burns. But neither design actually
accelerated at all as an airbreather.Scramjet is NOT an off-the-shelf ready-to-apply
technology. It just barely works at all, in a few
highly-experimental and hugely expensive flight tests. X-43A flew
twice out of 3 attempts, netting two 3-second burns at constant speeds of
Mach 7 and 10, with hydrogen fuel. X-51A flew twice out of four
attempts, netting two 3-minute burns at constant speeds of Mach 5,
with hydrocarbon fuel.

These scramjets can indeed
be developed, yes, but I wouldn't hold my breath waiting for a
ready-to-apply technology! I have personally watched this endeavor for
over half a century, without seeing anything but highly experimental,
only-partially successful results,
and only in the last 12 years. Nothing before that.

Acceleration
in Airbreather Mode

(Conventional, subsonic-combustion) ramjet is something I
did for a living for ~20 years in defense work. Most of these were
designed for max speeds of Mach 3 to Mach 4. Although, one test vehicle (ASALM-PTV) reached an
unintended Mach 6 on a short transient. It accelerated airbreather-only
from Mach 2.5 takeover to Mach 6, in a matter of several seconds! Vehicle
thrust minus drag, divided by vehicle
weight, was a fair fraction of one full
gee! Like I said above, ramjet
works, is off-the-shelf, and is very definitely effective and
ready-to-apply.

Combined-Cycle
Engines

Combined cycle engine ideas are like scramjet, just
experimental toys. Nothing is off-the-shelf ready-to-apply. I've also been watching that effort for
nearly half a century. The closest thing to reality is the ejector
ramjet, but you'll actually get better overall performance if you just
physically separate the rocket from the ramjet, which then offers some unique
advantages for parallel burn. Each can then assist the other anywhere
needed, all along the trajectory, and
this happens at the best (uncompromised) performance from each.

There are
fundamentally two kinds of combined-cycle engines: the turbine-based and rocket-based
combined-cycle engines. It would be hard
enough to do either of these as turbine- or rocket- combined with ramjet, but the trend in recent decades has been to try
to blend them with scramjet. This has
turned out to be utter nonsense so far, not
just because combined-cycle makes fatally-severe compromises to individual
component performance, but also because
scramjet is just fundamentally unready to apply, even in its pure forms.

The rocket-based
combined cycle with conventional subsonic-combustion ramjet is also known as
the ejector ramjet. This one hides a
rocket engine within the ramjet engine,
in an effort to provide meaningful static thrust for takeoff. It actually can work, but the performance levels are abysmally
low. It’s just not worth the losses.

The rocket’s thrust
performance is greatly reduced by the jet drag against the ramjet structures
around it. Any thrust-augmenting airflow
induced through the engine by the rocket jet is drastically reduced for all
inlets with external compression features,
simply because there are no attached-shock solutions for the external
compression features. That’s just
physics, there is no way around it. That effect inherently “kills” ingested air
flow. And the flameholding flow pattern
of the ramjet component is inherently and fatally disturbed by the presence of
the rocket jet, with which it is
coaxial.

The
concept of separate rocket and ramjet engines capable of “parallel burn” is just
the better deal, and by far! The
individual components yield full performance levels, and can be operated simultaneously as well as
sequentially, which is far greater
flexibility. Parallel burn with separate
devices is also how a self-boosting ramjet airplane becomes a feasible thing to
attempt. You need to retain enough
rocket propellant to provide a safe and practical landing capability: “go around” or “divert” on rocket
propulsion. This is probably best done
with small liquid propellant rocket engines to provide boost, landing,
and any parallel-burn mission capabilities. These can use the same fuel as the ramjet, so that only the oxidizer need be added.

Ramjet Heat Protection

Up to now, ramjet has been applied to one-shot missile
designs. Combustor heat protection is
best done as DC-93-104 silicone ablative,
but retained in place once charred-through, by kinked stainless steel ribbons spot-welded
to the case ID. This retained char
becomes brittle and falls apart after the burn,
upon cool-down, so it simply cannot
be reused! The kinked-ribbon
retention idea has been implemented by very few outfits!

However, by means of acid-etched Teflon film and
appropriate primers, conventional rocket
propellant can be cast in place on DC-93-104 as a fully-case-bonded integral
booster rocket. Without the Teflon
separator, this does not work, as the silicone chemistry of the DC-93-104 is
fundamentally incompatible with the hydrocarbon chemistry of CTPB, HTPB,
and PBAN propellant binder systems.
(I leave out monopropellant explosive GAP binder as simply too hazardous
to work with, like raw NG, but it would also need the Teflon barrier
film.) The most practical propellant grain
design for ramjet-like L/D’s and high boost pressures is the keyhole slot.

High-density
“super-ceramics” are just not a feasible option for a combustor packaged
inside an airframe, precisely because
they have high thermal conductivity (inherently because of their high density). The heat flow through such a wall is
catastrophically large, destroying
anything near the incandescent combustor shell.
The only “option” would be regenerative cooling, but unlike rocket, the fuel flow is far smaller and so unable to
accept the large quantity of heat. This
same consideration applies to nozzle and inlet structures.

There are a couple of
low-density ceramic solutions, but these
are very experimental, and preclude
the use of an integral booster packaged in the combustor. For surface temperatures under 3200 F, an all-alumino-silicate solution is available
as a ceramic composite made of pipe insulation paste and aircraft fire curtain
cloth. This has been done once
successfully, in a miniature combustor
whose flight speed did not exceed Mach 0.9.

For surface temperatures
exceeding 3200 F, some kind of
reduced-density zirconia composite is required.
The material choices are fewer,
and far more expensive. Thermal
conductivity is very much higher, so the
required thickness is far greater. As
near as I can tell, this was actually
attempted once, but has never been
done successfully. There are also
very serious concerns about usable lifetime,
on the part of the zirconia materials maker.

Neither
ceramic solution is compatible with the integral booster concept. All
such materials, being so porous, are inherently very fragile. Booster rocket pressures would utterly
destroy them in microseconds. For either
ceramic solution, the booster rocket must
be located outside the combustor,
which means either a stage-off booster or else a parallel-burn rocket of
some kind. There is simply no way around
that.

Ramjet missiles
usually feature exposed martensitic stainless-steel structures, good to about 1000-maybe-1200-F material
temperatures. Plain carbon steel and
titanium are only good to about 700-800 F material temperatures. That 1200 F limit with stuff like D6ac is why
no operational ramjet missiles have ever exceeded about Mach 4 cruise
velocity. To fly faster requires some
way to limit material temperatures. Air
temperatures in the stratosphere are around 3000 F at Mach 6. And that’s just skins. Leading edges are more demanding.

The only other ramjet
heat protection scheme is the perforated cooling air sleeve, something common in designs from the 1950’s
and 1960’s. It is still seen in jet
engine afterburner ducts. Overall
mixture must be lean to have the excess air needed for the cooling sleeve, and this air must be cool enough to actually
serve effectively as a coolant for an item washed by flame on the other
side. Such schemes have never served at
speeds over Mach 3 because of the air temperatures. A perforated sleeve like this is as
incompatible with an integral booster as are the low-density ceramics.

Ramjet Flame Stability

For all the
non-hypergolic fuels, there must be
properly sized and located flameholding recirculation zones, or else the igniter must fire massively throughout
the flight. There is no way around that, it is just physics and chemistry. The hypergolic fuels are vapor magnesium and
TEB or TEB/TEA blends. These are low
energy, low stoichiometry, low-Isp (very high TSFC) fuels. They are also the only feasible flight igniter
materials.

No flameholding ramjet
(gasoline, wide-cut, kerosene,
or solid gas generator-fed) was ever successful with a ramjet throat
area / combustor area ratio exceeding 65%,
because the balance of inlet and duct sizes becomes impossible at
feasible flow speeds above that limit.
Few were ever successful with that ratio below about 55-60%, because frontal thrust densities fell too low
to be useful. It makes sense to size at
65%, and set inlet/throat area ratio so
as not to spill, throughout the flight
envelope, and duct area so as not to be
unignitable or suffer flashback from high or low duct velocities.

Hypergolic (magnesium
vapor) gas generator-fed systems can use throat area ratios up to ~90%, and need no flameholder recirculation zones
at all. All that is important is
fine-scale mixing for combustion efficiency.
But hypergolic systems are the exception, not the rule.
The SA-6 was one of these. There
have been no others, with the possible
exception of the ramjet variant of the AA-12 “Adder”.

There are two practical
kinds of flameholder: baffles and “sudden-dump”
combustors. The actually-implemented
form of the baffle is the V-gutter, used
in the 1950’s and 1960’s systems, and
still in use today in jet engine afterburners.
These are very sensitive to the speed in the duct approaching the
V-gutter. They do require inlet air
temperatures low enough to be an effective coolant, because these structures are bathed in flame
on the downstream side. They have never
been successfully used as bare metal items above about Mach 3 or so, because of the high inlet air temperatures. Their area blockage is usually no more than
about 10-15% of the inlet duct area. These effects depend upon absolute
size: smaller is less stable.

The sudden dump
combustor, whether center or side
entry, has only insulated combustor
surfaces facing the combustor flame.
These can be used to at least Mach 6,
which is about the practical limit with subsonic-combustion ramjet
anyway. They are relatively insensitive
to approaching duct velocity, and thus
more stable for a wider range of conditions.
It is low pressure and low temperature in the oncoming air that have the
greatest negative impact on flame stability in dump combustors. The sudden
expansion area ratio must fall in the feasible range for this to work. These effects also depend upon absolute
size: smaller is less stable.

In ASALM-PTV, the fuel spraybar assembly was configured to
shed a wake resembling that of a V-gutter,
but it was insulated on the downstream side, and located right in the dump plane, in order to support the inlet port cover
during boost. The purpose of that
spraybar wake turbulence was additional mixing energy to offset the effects of
a very short combustor, not to enhance
flameholding.

The other practical baffle
flameholder (besides the V-gutter) is the “colander” burner, whether inverted or not. This was last successfully used in a ramjet
in Talos in the 1950’s, and that proved
to be fraught with problems in testing.
It finds wide application in gas turbine engine work, where flowing conditions are just not quite so
extreme. It does not find
application in afterburner ducts, which
are actually quite similar to ramjets.

Ramjet
Engine Sizing

Ramjet operation and
performance is almost completely determined by what the inlet can do. If the pressure demanded of the inlet is too
high, the shock system is expelled out in
front of the cowl lip, leading to
spillage of decelerated air. That spillage
incurs drag without the benefit of producing any thrust.

Things that raise the
pressure demanded of the inlet are (1) too large an inlet relative to the other
proportions, (2) too rich a
mixture, and (3) too small a ramjet
throat relative to the other proportions.
Inlet, throat, and mixture strength must be carefully
determined to meet requirements and also maximize performance while doing
so. This is not a trivial exercise; it requires both appropriate knowledge and
real experience.

The subsonic inlet
duct must also be sized in correct proportion,
so that it has an appropriate flow speed, both burning,
and in cold flow for ignition.
Flameholding is a part of this (upper speed limits), as is the risk of flashback (lower speed
limits), which can destroy inlet
structures. So also is the flame
stabilizer pressure loss a function of chosen duct speed. Most dump-stabilized systems do well if the ratio of duct
area to combustor area falls between 45% and 50%.

The only pressure-rise
item in a ramjet is the inlet recovery.
The combustor entry and flameholder are pressure losses, as is the combustion zone itself. This is quite unlike the gas turbine, in which the pressure rise effect of the compressor
dominates, and by far. This compressor pressure rise is always much
larger than the pressure drop across the turbine that drives the
compressor, in turn the dominant
pressure reduction effect.

Ramjet thrust and
performance maximize for maximum captured air massflow, with maximum feasible pressure recovery. You cannot truly maximize both
simultaneously, and massflow is the more
important item in ramjet. Therefore, one operates with a critical-to-slightly-supercritical
inlet for max air ingestion at good pressure recovery. This is quite distinct from gas turbine, which requires spillage to match inlet and
engine massflows: an always-subcritical
inlet. The inlets of ramjets and gas
turbines look the same (in the same speed ranges) because they utilize the same
components, but these are actually used
quite differently.

Ramjet
Ground Testing

Testing ramjets on the
ground is best done in a direct-connect test facility, as long as the inlet performance is already
well-defined. Most of these use
“vitiated” (combustion-heated) air with make-up oxygen to achieve the needed temperatures.
This is appropriate for the liquid hydrocarbon fuels, but not the metallized gas generator
fuels. The active metals utilize the
excess carbon dioxide and water vapor in vitiated air as additional oxygen, thus getting bad test results. Similar problems increasingly cloud results
even with the liquid fuels at inlet air temperatures beyond about 2500 F, due to dissociation and ionization chemistry
effects.

One overcomes this by
means of a pebble bed-type “clean air” heater instead of combustion
vitiation. Achievable temperatures are somewhat
lower with this design approach, but the
delivered air really is air. This is
crucial with the metallic hypergolic fuels,
like magnesium. Even aluminum and
boron are questionable.

Analysis of ground
test data is not done by comparison to predictions from a cycle code. Instead,
many of the same mathematical models are incorporated into a test data
analysis program that computes independent estimates of combustion performance
from both combustor pressure and test article thrust, as completely-separate sources of data. When these agree, you know you did everything right. Of the two,
getting reliable thrust data is far more difficult, because of the difficulty of calibrating all
the possible facility tare forces. There
is no such thing as a tare pressure,
though. Trust pressure-derived
performance in preference to thrust-derived performance, always.

Subscale
Test Scaling

Scaling down to
subscale test is more complicated than generally thought. What you want is the same pressure and speed
distributions inside the subscale representation of your engine. If no other considerations were
important, then all you need do is match
the air and fuel flow rates per unit cross section area, the inlet total temperature, and the geometric proportions of the
engine. But, other considerations do matter. Once you scale down too far, the residence time distribution will be
fatally wrong compared to full scale!
This is because chemistry rates do not scale with size.

There is a minimum
size below which you have to distort the engine geometry in order to maintain
feasible residence time distributions.
Only certain kinds of distortions are effective. The details of this are entirely different
for each geometry class (baffle versus coaxial dump versus side entry dump). There
is no one general procedure to use! The
criteria are entirely empirical and unique in each geometry class. Both overall residence time and flameholder
recirculation zone residence time are critical items to address. These are computed quite differently, and in the case of flameholder residence
time, methods differ by geometry class.

Running
Studies Requires Modeling Inlets

A lot of the folks who
want to do this, do not have real
wind tunnel data on real inlets available to them. As long as the studies are “ballpark”
explorations, and not real system
predictions, there is a way to adequately
estimate inlet performance for “new designs” based on past historical
data. This estimating technique is based
on the “shock-on-lip design Mach number” of the inlet, which applies to high-speed inlet designs
only.

I have a curve
composited from old data that is fairly universal, if used as the pressure recovery PRCR for
Mach numbers below shock-on-lip. In that
regime, they’re basically all just about
the same. A factor taken from a second
curve applies to the PRCR value at shock-on-lip, to create the individualized PRCR curve at
Mach numbers above shock-on-lip. That
factor varies with flight-minus-design difference in Mach numbers.

I have two other
curves that model streamtube area ratio ARCR trends with that same difference
of flight minus design Mach. One is for
round inlet cross-section shapes, the
other is for two-dimensional inlet cross-section shapes. The basis area is AC, defined as the swept-out area: this is the cross-section of the cowl entry
channel plus the frontal blockage of the external compression surfaces, at zero angle of attack.

These techniques work
fairly well for near-zero angle of attack with side-mounted inlets and chin
inlets, and pretty much up to 15 or 20
degree angle of attack with nose inlets.
It is easy to use this technique in a trade study to help define what
the “best” inlet shock-on-lip design speed is, for any given problem.

The missing piece is
additive drag coefficient, which does
not usually apply to nose inlets at all,
may or may not apply to a chin inlet,
but is quite important for side-mounted inlets. This usually represents the ”pre-entry” drag
on the entering streamtube, where it is
in contact with vehicle surfaces and influenced by the vehicle bow and forebody
shock and expansion field. It does not
include the spillage drag when operating subcritically. The reference area for this is also AC, as defined above. This coefficient, the AC, and the freestream dynamic pressure multiply
together for the additive drag force.

Often, the effects of boundary layer diverter
drag, and the spillage drag of capture
enhancing bleeds located near the cowl lip,
are included lumped-in with the pre-entry drag into the additive drag
data. I have a set of real wind tunnel
additive drag data that includes pre-entry,
diverter, and capture-enhancing
bleed drags for a real design actually tested. There is a knee in this curve at
the design shock-on-lip speed. For trade
study purposes, I just shift this curve
left or right to put that knee at the shock-on-lip speed in my study
problem.

Subcritical spillage
drag coefficient is easy to estimate as twice the subcritical spillage
margin. The area basis is AC, and those with the free stream dynamic
pressure gets you to a drag force for the subcritically-spilled air.

Many systems use air
bled from the subsonic diffuser aft of the terminal shock, to power pneumatically-operated
machinery. This bleed reduces the air
actually fed to combustor from that captured by the inlet, by an amount called the “bleed
fraction”. As long as your ram drag
(inlet air momentum) is based on all the air captured, you have already accounted for the drag of
scooping up your machinery bleed air.

Pitot/normal shock
inlets for low-speed designs are far easier to estimate. These are almost invariably nose inlets
without any capture-enhancing bleeds, so
the additive drag is zero.

Pick a “high” number
like 98% to represent the subsonic diffuser PRCR. From Mach one on up, multiply that 98% factor by the total
pressure ratio across a normal shock at each Mach number. That product is your supersonic PRCR, rather closely.

Your pitot/normal
shock ARCR is just another pretty-constant “high” factor like 98%, across the board from subsonic to supersonic.
Base your AC on the dividing-streamline “highlight” defined by the
inlet lip radius.

Thrust-Drag
Accounting

There are two
systems: (1) net jet, and (2) installed. As regards ramjet propulsion, net jet thrust
is nozzle thrust minus the ram drag of the captured airflow. The additive and subcritical spillage drags
must be added to the airframe drag.
Installed thrust is nozzle thrust minus ram drag, minus additive drag, and minus subcritical spillage drag. The airframe drag is unchanged. Do not mix definitions! Be consistent! Net jet is popular among propulsion
specialists, while the vehicle
aerodynamicists and trajectory dynamics folks prefer installed.

Engine
Flight Envelope

The standard
presentation is altitude on the vertical axis,
and speed on the horizontal axis.
For supersonic-capable systems of all kinds, the preferred form for speed is Mach
number. You can create one such plot for
each day-type model (such as a “standard day”) that you choose to use.

The minimum ramjet
operating speed on this plot may or may not be constant with altitude. Generally it is not the absolute minimum operating
speed for thrust equal to drag, but
something higher set by adequate vehicle acceleration capability. The maximum ramjet operating speed is very
likely to be determined by thermal protection risks. Something like 1200 F inlet air total
temperature is usually a good representation of this.

The “ceiling” of this
operating envelope is fairly likely to be either a scooped air massflow
contour, or a flight dynamic pressure
contour. Flameholder stability and fuel
turndown ratio limitations may clip off corners or zones from this basic
envelope, as your study proceeds. The process of drawing it is iterative.

A variation on this
uses the scaled-down inlet AC to generate contours of constant
airflow on a flight envelope that corresponds to subscale test hardware. These and the constant inlet temperature
contours get plotted as an easy way to relate ground test conditions (airflow
and temperature) to simulated flight conditions (Mach number and altitude) for
that particular design. Facility limits
usually exist in the form of air flow rate and temperature limits for direct
comparison. Open-air nozzle choke limits
can usually be determined in terms of a critical airflow value. Where you can simulate in test is thus a
sub-envelope, generally.

Selecting
Propulsion

Gas turbine is
available basically in two forms: (1)
high-bypass ratio “fanjets” that offer high economy, but only at subsonic speeds, and (2) low bypass ratio “turbojets” that
offer supersonic flight at the cost of substantially-lower fuel economy. Both types can be thrust-augmented with
afterburners, at the cost of very low
economy. In practical terms, gas turbine
is limited by excessive inlet air temperatures to maximum feasible speeds near
Mach 3.3-3.5 in the stratosphere.

High-speed range
ramjets are useful from near Mach 2, to
at least Mach 4, and perhaps to Mach
6. They are far simpler and lighter than
gas turbines, but resemble afterburning low-bypass
gas turbines in terms of fuel economy.
Simple, lightweight, inexpensive,
and more fuel-economical than solid rocket propulsion, that is why ramjets are often employed as
missile propulsion.

Ramjets are not as
simple to design, or to incorporate into
a compact missile, as solid rockets, so they are not generally selected for the
shorter range tactical missiles. For the
longer stand-off range tactical missions,
ramjets are well worth the trouble to incorporate, especially the modern integral-booster
forms. Ramjets will cover the range in
smaller packages, arriving at higher
speeds for better maneuverability, and will
do so in shorter flight times that greatly enhance the survivability of the
launch aircraft, ship, or site.

Once there really are
ready-to-apply scramjet and combined-cycle engine technologies, you will have more options to choose
from. That time is not yet.

Concluding
Comments

The details of all
items just discussed were obviously not included. There are many more details and issues
associated with ramjet-propelled vehicles,
all quite critical to success.
The sum of all that is far larger than can be put into a few paragraphs
here.

My ramjet book is
still in work, but it is more than
half-written now. It addresses all these
issues, and much more besides, in a very hands-on / how-to manner. In it,
I tried to include not only the science,
the applications, real
examples, and the history, but also a lot of the engineering art of ramjet
propulsion that I do happen to know.
That art is the part not written down,
but passed-on in the workplace directly from the seasoned hand to the
newbie.

I hope to find the
proper publishing outlet, get this book
finished, and get it published and
available, during 2017.

As regards engineering
art, I am fond of saying that “rocket
science ain’t science, it’s only about
40% science. It’s about 50% art, and about 10% blind dumb luck”. I would also add two extra points to that
statement:

(1) It applies to
production work. In development
work, the art and luck factors are even
higher.

(2) It applies to just
about all of engineering, not just
rocketry or ramjetting.

Sunday, October 2, 2016

Every year, the
International Astronautical Federation (IAF),
the International Academy of Astronautics (IAA), and the International Institute of Space Law
(IISL) hold a meeting somewhere in the world called the International
Astronautical Congress (IAC). The latest
one was September 26-30 in Guadalajara,
Mexico. The next ones are:

Elon Musk of Spacex gave a very astonishing presentation at
that recent meeting, revealing how he
plans to enable a settlement on Mars at an affordable price, sooner than anyone
ever thought. I saw the on-line video of
his presentation, and downloaded the
slides he used.

After taking some time to digest both content, and the comment and criticism this
presentation has generated, here is my
best shot at the most-factual summary.
This is not a full-blown colony that Mr. Musk proposes to establish all
by himself, despite what the news story
headlines say.

He is creating the essential practical transportation system
that is necessary to enable many entities to participate in establishing a permanent
colony. No one entity can do this. The timeline for creating such a colony is
around a century long.

Some of the comments I have seen claim Musk is living in
“fantasyland”, because his proposals are
so vastly different than anything we have ever seen out of NASA. They look almost like the 1940’s-1950’s
dreams of big spaceships that don’t seem limited in their range.

But I disagree with the critics; Musk can now take advantage of enabling
technologies and materials today, that
were simply unavailable to NASA or anyone else in prior decades. If he does embark on this plan, and he can successfully pull off developing
these vehicles and flying them, it will
fundamentally and forever change the world of human space travel.

Can he do this? His
history says that he does what he says he will do, but just not quite as fast as he wanted to do
it, because space flight is hard, with many setbacks. As proof,
I cite the development of Falcon family of launch vehicles that has so
very dramatically helped to reduce the cost of expendable launch to Earth
orbit. Musk has already begun to recover
those boosters, and will soon attempt re-flying
them (a key feature of the Mars ship he proposes).

This isn’t something Musk and Spacex can do all alone. They have to make money by selling rides in
this giant vehicle. This thing offers
something like 300 ton capacity to low Earth orbit, and at a price per delivered ton far below
today’s prices. This completely opens up
the field for entrepreneurs wanting to build space stations in Earth orbit for
all sorts of purposes.

Musk’s “big rocket” is really two things, a giant launch booster based on scaled-up
Falcon technology, and a spacecraft that
is also its second stage getting to orbit.
The booster flies back to launch site for reuse, and the second stage/spaceship is refueled in
orbit to enable it to fly to any desired low-gravity destination, without further staging.

Musk's 2-Stage Vehicle with 2nd Stage the Interplanetary Spaceship

Data on the Reusable 1st Stage Booster

Data on the Reusable 2nd Stage Spaceship (there is a tanker version)

Depiction of Booster Fly-Back Similar to Falcon Boosters

There are actually two versions of that second
stage/spaceship: one is the spaceship
that must be refueled in orbit to go anywhere else. The other version is a tanker that does the
refueling in orbit, and then flies home
for reuse. But, for the spaceship to come home from its trip, its crew must make more propellant at that
destination.

This thing dwarfs the old Saturn 5 moon rocket. At launch it’s about 400 feet tall, 3.5 times the weight, and 3.6 times the thrust, of the Saturn-5. It also dwarfs NASA’s three-stage Space
Launch System (SLS) rocket, currently in
development, and which is not reusable
at all.

Comparison of Mars Vehicle and Booster with Saturn-5

There are lots of low-gravity places that a vehicle like
this can go. These include Mars, our moon,
any moon of Jupiter or Saturn, and
the asteroids and comets. Maybe
Mercury. Maybe the outer planet
moons. But not high-gravity places like Earth
without a booster and tankers. (Venus
isn’t a feasible destination for several reasons besides its gravity.)

Musk’s schedule calls for development and testing of actual
hardware beginning 2018, aimed at a first
Mars flight late in 2022.
Personally, I think it’ll take a
little longer than that, precisely because
space travel is hard and there will be setbacks. There already have been, with his Falcon rockets. But I would guess his Mars ship, tankers, and giant booster will be flying no later than
about 2030, maybe a little sooner.

His presentation showed launches out of Cape Canaveral. But I think he really wants to launch this thing out of his private space launch facility in deep south Texas.

Location of South Texas Private Launch Site

This thing has enough performance to fly a little faster to
Mars than an absolute minimum-energy trajectory. That shortens trip times considerably, from 8 months to 3-4 months, one-way.

Shorter flight times mean we are “OK” with no artificial
gravity and no high-efficiency life support systems, and that the accumulated space radiation dose
is much lower. There’s plenty of room
inside the ship for maybe a hundred people to live, for journeys of that shorter duration. Only for further destinations would these
issues need to be addressed.

Once at Mars, the
ship makes a direct entry from its interplanetary trajectory and a
retropropulsive landing without parachutes.
Direct entry is something NASA has done for decades with its
probes. Doing propulsive landings
without parachutes is something NASA has never done, except on the moon with Apollo, and never since. Problem is,
chutes are ineffective on Mars for masses over about a ton.

This Mars mission architecture depends upon making
propellant on Mars, because the vehicle
uses up all its propellants getting there.
In that sense, Musk is betting
the lives of the ship’s occupants that they can make enough propellants on Mars
fast enough to refuel it for the journey home.
It needs no booster to do that:
Mars gravity is only 38% that of Earth.
The ship holds over 1900 tons of propellants, as currently envisioned. That’s a lot to make in only a few months!

The choice of which propellants to make is crucial for
success. Musk and Spacex have chosen
liquid methane and liquid oxygen, made
with the local Martian “air” (which is 98% carbon dioxide), and local water-as-ice. There is a chemical process called “Sabatier”
that uses carbon dioxide, water, and electricity, to make methane and oxygen. Then you liquefy them, which also requires electricity.

The key to this is finding local sources of ice to melt for
the water.It means the astronauts are
going to have to be ice miners when they get to Mars. The nature of the buried ice deposits will
determine the complexity and weight of the tools, equipment,
and facilities that the vehicle must transport there.

If there is a massive buried glacier at the landing
site, then one can simply slant-well
drill into it, use hot steam from a
steam generator to melt the subsurface ice,
and bring gobs of water right back up the well. If the ice veins are thin layers, or separated pockets, strip mining will instead be required, with enormous dirt-moving equipment, and a gigantic facility processing enormous
volumes of surface material, for each
precious ton of water.

Obviously, selecting
the right landing site is critical.

Musk’s plans begin with what he already has in-hand
today. There is a new version of the
Falcon called Falcon-Heavy, which should
start flying in 2017. This rocket is
powerful enough to send a version of his Dragon capsule unmanned to Mars, for a direct retropropulsive landing, with 2+ tons of “stuff” on board.

That version is called “Red Dragon”, and will carry as-yet unidentified robotic
payloads to investigate multiple potential landing sites. Musk expects to start these flights in
2018. NASA is finally participating, hoping to learn about retropropulsive landings
on Mars, but is too late for them to add any
payload items to that first flight.

Depiction of Unmanned Red Dragon Landing on Mars

So, what would
constitute the right landing site?

First, Musk’s Mars
ship is 3-4 times as tall as its landing legs are wide. That makes it very intolerant of rough ground,
or obstructions like boulders and
dunes. The site must be very flat and
clean of hazards.

Second, there needs
to be massive buried ice deposits directly underneath the landing site. That’s something remote sensing is just not
capable of determining. Ground truth has
always been at variance with remote-sensing claims, often enormously so.

It will take real drilling to determine this,
just like it does here exploring for water or for oil/gas/coal. Whatever payloads Red Dragon carries, a robot drill rig capable of drilling at
least a football-field down is required.

A Canadian outfit called NORCAT built a robot drill rig that
it called “CanaDrill” a few years ago.
They offered it to NASA without success,
my sources tell me. But, to me,
it looks like some version of this thing is exactly what Musk needs to
ride his Red Dragons to Mars, looking
for that “right” landing site.

So how will this “play out”?

Falcon-Heavy/Red Dragon shots identify the right landing
site between 2018 and the time the big Mars ship is ready to make its first
Mars flight.

The first big Mars ship flight brings mostly cargo and a
small crew. This crew sets up that first
“outpost” as a modest habitat in which to live, and a minimally-adequate propellant-making
plant for the return trip. They do human
exploration of Mars while there, of
course. As the flights continue every
two years when the orbits are favorable,
the propellant plant grows in capacity,
and the habitat also starts enlarging to accommodate larger populations
later.

Based on our history attempting such things so far, I have very serious doubts that a
closed-cycle ecology, a self-sustaining
life support and food production scheme,
is going to be successful in the early years of this outpost. That means it will initially be dependent
upon regular resupply, as part of the
cargoes of these big ships.

Later on, this issue
will get resolved, and the outpost
becomes essentially self-supporting from a life support/food production
standpoint. That’s the point at which it
can first evolve into a real “city on Mars”,
with some serious local production of supplies and infrastructure items
(yet to be identified). By that
time, the ships will be bringing more
people than equipment and supplies.

Eventually, somewhere
in this process, some sort of Martian
exports yet-to-be-identified will be making the return voyages to Earth. Some sort of interplanetary economy will
evolve from that. That is the point at
which you can really call this a proper “colony on Mars”. I think (as does Musk and Spacex) that’s
about a century down the road.

Comparison to NASA/”Big Space” business-as-usual:

Musk wants people on the surface of Mars by about 2025.Realistically, I think he will do this closer to 2030. Compare that to NASA/”Big Space’s” plans to
fly around Mars without landing in the late 2030’s, with the actual first landing sometime in the
2040’s. Musk beats them by at least a
decade.

Musk has already begun to develop his fully-reusable
spaceships of enormous cargo capacity.
He has already started landing boosters,
he will soon re-fly used boosters,
he already has a good heat shield for entry at Mars or returning to
Earth, he is already starting to test
his methane-oxygen rocket engine (see photo), and he is already constructing his first giant
propellant tank test articles needed for the big booster and ship. Compare that to the cramped capsules and
throw-away stages seen proposed by NASA/”Big Space”, or by any another entity on the planet.

Raptor Engine Test (Big Methane-Oxygen Engine)

According to Musk’s presentation, they are projecting around $200 million per
launch of their reusable boosters and ships.
They project a price under $200,000 for each ton delivered to Mars, which is also roughly the same as the ticket
price per person. Compare that to around
$1 billion per launch of NASA’s SLS, and
a per-astronaut cost to Mars in the 10’s of billions of dollars.

What makes this affordable transportation possible at
all, and what makes this plan look like
something out of the dreams from the 1940’s and 1950’s, is this specific list of enabling
technologies, taken right from Musk’s
own slides:

Refueling
in orbit (NASA doesn’t do this; the
Russians do, but not with cryogenic
propellants)

Select
the right propellant that can be manufactured at destination (NASA doesn’t do
this, yet)

Manufacture
of said propellant at destination (NASA doesn’t do this, yet)

This is a total “leap-frog jump” into the future, that leaves everyone else behind. I think it is really possible to do
this, although I also think it will be
harder to accomplish than it looks to Musk and Spacex right now.
Being a life-long fan of human space travel, I can only wish them success.

Update 10-8-16: There should probably be a 5th item in Mr. Musk's list of 4 enabling items just above: gigantic size. He and Spacex have thrown the artificial, self-imposed "minimum thrown weight" constraint right out the window. This is unlike anything proposed since the giant spaceship concepts of the mid-1950's. The New World was not settled from Europe with small boats. They used the full-size ships of that time. The airline industry in the US was not started successfully with small airplanes that had only a few passenger seats. It took Ford leaping in with the Tri-Motor, and Douglas leaping in with the DC-3, to point the right way: that large aircraft were what really worked. The same is true here: one of the cost savers with Musk's giant rocket is simple economy-of-scale. Update 10-28-16: Do not be confused by the talk of 100 people to Mars per ship, or colonies with a million people. That's "far future" stuff, some decades after the initial landings, if not longer time spans. Musk is concentrating on the transportation system, not the actual establishment of a colony. He cannot do all of this by himself. He will be very lucky just to get the transportation system done. There is a timeline disparity in what he presented at Guadalajara: first giant ship to Mars late 2022, versus first men to Mars late 2024. This is preceded by Falcon-Heavy/Red Dragon "pathfinder" unmanned shots in 2018, 2020, and presumably into 2022. To me, this sort-of looks like the first large-ship shot or two is unmanned. Presumably, that would have something to do with the return propellant processing factory, and perhaps other infrastructure, for a putative manned base to start around 2022. And, as always, calibrate this with the history so far. Musk typically does what he says he will do; it just takes him about 50+% longer than he wanted, to actually get it done.

----------------------------

If you go to "Mars Mission Outline 2016" on this site, dated 5-28-2016, you can see what I had been working on. Musk and I share the concepts of big ships, and of re-usability to the maximum extent possible. Where we differed was (1) I used separate landers rather than a direct landing, and (2) I did not presume local propellant manufacturing capability from the very first manned landing.

That is why the two transportation systems look so very different. Assumptions make a gigantic difference.

Followers

About Me

"GW" is an aerospace engineer and a teacher, with considerable experience in both professions. He is currently looking for a little consulting, but is now retired.
............................................................
He has done both automotive and aircraft alternate fuel work. The photo is of GW in his "ethanol VW", which runs on straight E-85. It is now mothballed, but his antique farm tractor runs on E-85, and just about everything else on E-30 to E-35 blend.
............................................................
GW sells, and now rents out, cactus-killing farm implements. He builds these in his shop out on the Texas Idea Farm.